For large aircraft, a wake of rotating air is created behind the craft during flight. These vortices and their associated downflow (airflow located between the vortices) and upflow (airflow located outside the vortices) can disrupt the flight of a following aircraft that enters the wake of the leading aircraft. Depending on the conditions, structural stresses can be imposed on the following aircraft. In some instances, the following aircraft can be rapidly rolled to one side, which is particularly hazardous when flying close to the ground. The problems associated with wake vortices become more pronounced as the lead aircraft increases in size relative to the following aircraft.
While the trailing vortices are almost always moved or destroyed by atmospheric currents, or spontaneously descend away from the flight paths, these effects do not guarantee that hazardous encounters will never occur. Several accidents have been attributed to wake turbulence. To alleviate this risk, a number of regulatory flight rules are in effect which require compliance in the amount of spacing between airplanes in approach to the same airport. The wake turbulence spacing requirement is often larger than the spacing required by other factors such as radar resolution or runway occupancy, so that under some conditions, these rules add to the congestion and delays in the air transportation system.
Starting around 1970, theories regarding the alleviation of wake vortices have been the subject of a number of professional conferences. The inventors herein are not aware of any such theory having been applied directly to a system for use with commercial airliners. Prior attempts to manage the trailing vortices in the commercial aircraft industry have focused on methods of modifying the vortex characteristics to alleviate (or reduce) the potential for upset on the following aircraft. Methods of this type provide a modification of the vortex structure, but such modification does not significantly alter the induced rolling moments on a following aircraft unless the following aircraft wing span is comparable to the vortex-core diameter. Typically, the vortex-core diameter is much smaller than the leading aircraft wing span. This type of system has been the subject of a large number of patents. Some of these are summarized in U.S. Pat. No. 5,492,289.
One of the dominant mechanisms affecting the persistence of the vortices is instability. Methods to excite a sinusoidal instability for a single pair of trailing vortices that could lead to a breakup of the vortices have been suggested. The growth of such instabilities would ultimately lead to large vortex distortions and eventual breakup. This idea was first analyzed by S. C. Crow in Crow, S. C., Stability Theory for a Pair of Trailing Vortices, AIAA Journal, 8, pp. 2172-2179 (1970) [hereinafter referred to as Crow (1970)]. A system of this type was suggested by Crow as documented in Crow, S. C, Panel Discussion in: "Aircraft Wake Turbulence and Its Detection," ed. J. Olsen, A. Goldburg, M. Rogers, pp. 551-582 (1971) [hereinafter referred to as Crow (1971)].
The predictions of Crow as to the most naturally-amplified wavelengths are in good agreement with experimental results where the natural source of excitation is provided by turbulence in the atmosphere. This source of breakup is referred to as the Crow Instability, and is well-known in the art. The Crow Instability amplifies slowly. The growth of the instability leads to periodic linking of the vortices from the right and left hand sides, in the flight-path direction in the far field. The linking transforms the vortices into vortex rings, which after some additional time, become sufficiently deformed to break up. In this final state of evolution, the impact of the vortices on a following aircraft can be considered benign.
The concept of directly exciting the Crow Instability was tested in towing-tank experiments reported in Bilanin, A. J. & Widnall, S. E., Aircraft Wake Dissipation By Sinusoidal Instability and Vortex Breakdown, AIAA Paper No. 73-107 (1973) [hereinafter referred to as Bilanin & Widnall]. The experiments showed that the instability could, in fact, be excited by varying the distribution of lift on the wing. A fuller presentation of the concept of direct excitation of the Crow Instability is given in the paper Crow, S. C. & Bate, E. R. Jr. 1976 Lifespan of Trailing Vortices In a Turbulent Atmosphere, J. Aircraft, 13, pp. 476-482 (1976) [hereinafter referred to as Crow & Bate].
The work of Crow & Bate represents the state-of-the-art of methods to excite a sinusoidal instability at the time of development of the current invention. It was suggested that the excitation would be accomplished by varying the distribution of lift on the wing in a particular manner. As stated in Crow & Bate, "The aircraft controls are wired so that the outboard ailerons can oscillate symmetrically, whereas the inboard ailerons move in opposition to keep the total lift of the wing constant. -The horizontal stabilizer trims any pitching tendency, and the net effect is to slosh the lift distribution in and out along the wingspan." Id. at pages 481. This concept is illustrated by FIGS. 1A and 1B herein labeled "Prior Art" and is taken directly from figure 10 of the Crow & Bate paper. FIG. 1A is a schematic of the control surfaces in their deflected positions. FIG. 1B shows a schematic of the wing spanload lift variation as suggested by Crow & Bate (the fuselage is not shown, but would be positioned at the center location of the x-axis.) The extreme of sloshing the lift outboard is given by the solid line, and the extreme of sloshing the lift inboard is given by the dashed line. The areas represented under the solid and dashed lines is the same, indicating that the two spanloads provide the same total airplane lift. However, the trailing vortices do not have the same strength (circulation) and location.
The centers of wake vorticity, called "vorticity centroids", are shown in FIG. 1C, where the dashed and solid circles correspond to the dashed and solid lines of FIG. 1B. Vorticity centroids are the centers of rotation for the combined vortices of one wing side, i.e., left and right vorticity centroids are shown in FIG.1C. Thus, FIGS. 1A, 1B, and 1C demonstrate that an inboard perturbation of a single vortex pair may be introduced by shifting the spanload lift (and hence the vorticity centroids) between the solid and dashed lines. In FIG. 1C, the circles represent the positions of each vortex and the lines represent the vorticity centroids on that side of the aircraft. Because the Crow Instability is theorized for only a single vortex pair, the centroids align with their respective vortex.
Referring to FIG. 1D, the vortices resulting from oscillating the lift distribution take on a sinusoidal form. The wavelength of the imposed perturbation is generally about seven wing spans (for maximum amplification due to the Crow Instability.) The vortex movements shown in FIG. 1D have been exaggerated.
As shown in FIG. 1C, significant movement of the vorticity centroids between the dashed and solid line positions for both the right and left hand sides is required to break up the vortices via direct excitation of the Crow Instability. To preserve the total lift while moving the vorticity centroids, the wing-root loading must also undergo significant variation, since the lift required from the wing-root is now much greater (see lift "hump" at the middle of the x-axis). Increased loading at the wing tip is also structurally problematic for many aircraft. Any necessary structural enhancements can lead to increased weight and fuel consumption which lower the efficiency of the craft.
The size of the initial vortex perturbation is a function of the amount of lift that is shifted along the wingspan. Thus, a large initial perturbation results in a large loss in lift from the baseline (steady-state) wing configuration. In order to break up the vortices within a distance behind the aircraft that could influence aircraft spacing, using the current approach (i.e., the approach of Crow & Bate) would require a large initial perturbation and an excessive loss in baseline wing lift. Because of the slow growth of the Crow Instability, the required excitation levels for rapid breakup of the vortices are too large to be useful for practical application. Indeed, the concept of Crow & Bate was never applied in practice.
Thus, a need exists for a practicable method and system or device for bringing about the dissipation of wake vortices by inserting a perturbation in the wake vortices that eventually causes the effective breakup of the vortices. The present invention is directed to fulfilling this need.